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25THINTERNATIONAL CONGRESS OF THE AERONAUTICAL SCIENCES

SUBSYSTEM DESIGN AND INTEGRATION FOR THE

MORE ELECTRIC AIRCRAFT

David Blanding

Technical FellowBoeing Phantom Works

Huntington Beach, CA

Keywords: More Electric, Actuator, Thermal, Fatigue, Aircraft

Abstract

Military and commercial aircraft designers are

leading a quiet revolution in the aviation

industry. Their goal is an all-electric aircraft

that will be controlled by small high speed

motors instead of heavy maintenance intensive

hydraulic, pneumatic and mechanical systems.

This revolutionary usage of electrical power

technologies promise military and commercial

airframers greater aircraft reliability and a

significantly smaller logistical tail to support

tomorrow's air and space force. Hence, the

More Electric Aircraft (MEA) is becoming a

reality.

The MEA approach provides for greater

integration of subsystem functions. It also

provides smarter subsystems without added

weight penalties. The MEA approach has

created common threads that link most or all

subsystems. These threads are:

Electrical power and distributionsystem

Thermal management system Integrated health management

The focus of this talk will address the

integration issues associated with the linkage

between electric actuation and the common

threads mentioned above. This paper will

focus specifically on issues associated with the

electromechanical actuator, address the

progress of the MEA, while at the same time,

discusses the growing popularity of advanced

materials that are enabling the MEA.

1.0 General Introduction)

Aircraft designers have already reached a

milestone installing MEA technologies in

airplanes like the Boeing 787 and various

military platforms. These systems represent the

first major leap in casting the MEA revolution.

They involve the revolutionary application of

electrical power systems, electronics, and

distributed architectures to simplify much of the

existing immensity bulk and complexity

inherent in traditional hydraulic and pneumatic

aircraft systems.

Emphasis now is on giving aircraft designers

more optional opportunity of using electrical

power over traditional methods. New

technology concepts like electric actuation,

electric environmental control and electric fuel

pumps, along with magnetic bearings for

generators and eventually more electric turbine

engines, are in the works. These technologies

promise dramatic simplifications in aircraft

system design, while improving reliability and

maintainability in the years to come.

1

Boeing Commercial Airplane (BCA) has chosen

to exploit the use of MEA technologies to

provide cost-efficient next generation airplanes.

The 787 is Boeings first More Electric

Airplane. Elimination of pneumatic system led

to No-Bleed architecture for overall airplane

weight savings and efficiency improvement.

Many airplane systems to are become

electrically powered for the first time. Some

examples are the 787 wing de-icing, largehydraulic pumps (used to raise landing gear),

flight control actuators (secondary systems),

Copyright 2006 The Boeing Company. All rights reserved

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David Blanding

cabin pressurization system, braking system,

and engine starting system. Advancement of

efficient and reliable power electronics

technology has enabled more electricarchitecture [1].

The MEA approach offers an increase in design

flexibility, a reduction in operation and

maintenance cost, and overall reduction in

system weight. A more notable benefit of the

MEA approach is the reduction in power

conversion, where you no longer have to

convert engine shaft power to electric, hydraulic

and pneumatic power (Figure 1). The extractionof single electrical power provides a cost

effective way to drive actuators, environmental

controls, fuel pumps, brakes and de-icing

systems [2]. See Figure 2.

Figure 1. Conventional Aircraft Power

Conversion

Figure 2. More-Electric/All-Electric Aircraft

Power Conversion

Electric Actuation denotes a broad applicationof electrical power to the actuation and control

of vehicle subsystems that traditionally have

been controlled by hydraulic, mechanical and

pneumatic power. One of the major drawbacks

for using electric actuators on primary flightcontrol surfaces has been the low power density

of traditional motors and thermally efficient

power electronics. There are three basic

approaches to electric actuation being developed

today for air vehicle flight control surfaces are

electromechanical actuators (EMA),

electrohydrostatic actuators (EHA), and

integrated actuator packages [2]. These electric

actuation systems are shown in Figure 3.

Figure 3. General Electric Actuation Technologies

For the purpose of this paper, the following

definitions of an electric actuation system apply:

Electromechanical Actuator (EMA) is

defined as an electric actuation system that

uses an electric motor mechanically

coupled to the load.

Electrohydrostatic Actuator (EHA) is

defined as a self-contained electricactuation system that uses an electric

motor driving a bi-directional, fixed

displacement hydraulic pump that operates

a typical hydraulic ram.

Integrated Actuator Package (IAP) is

defined as a self-contained electric

actuation system that uses an electric

motor driving a servo-over center

hydraulic pump that operates a typical

hydraulic ram.

2Copyright 2006 The Boeing Company. All rights reserved

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SUBSYSTEM DESIGN AND INTEGRATION FOR THE MORE ELECTRIC AIRCRAFT

Flight test experiments have been performed on

EMA, EHA and IAP systems to establish the

credibility of electric actuation as a primary

method of control for flight-critical controlsurfaces on tactical aircraft.

The EHA and IAP, which have been the

subjects of much industry and government

investments over the last decade, offer the

opportunity to produce the actuation power

electrically, using power-by-wire (PBW), while

retaining a jam-proof connection of the actuator

to the surface. The IAP has successfully

completed 1,000 operational hours using an

electric aileron actuation system, demonstrating

the reliability and maintainability benefits with

this technology. The EMA and EHA have also

been flight tested on various military platforms

to mature the technology.

In recent years the EMA enjoyed similar

attention on several vehicle development

programs. The EMA became the baseline

technology on these programs largely because

actuation requirements for these favored theEMA technology.

2. 0 Why Electromechanical Actuators?

EMA, EHA and IAP have their preferred

applications. Program managers and system

designers will continue to evaluate all three

technologies; however, if all three technologies

have the attributes for a particular application,

the EMA may be the preferred solution (Figure3) due to the absence of hydraulic fluid. This is

especially true for spacecraft applications and

vehicles with long storage requirements where

hydraulic fluid leakage and freezing may be a

concern. For the most part, an EHA and IAP

could be considered as similar technologies

because of the common use of a hydraulic ram

and pump.

Electromechanical actuators are a cost-effective

alternative because there is one energy

conversion versus two in a hydraulic system.

On some aerospace platforms, depending on the

size, analyses have shown that this technology

could provide weight savings averaging from afew hundreds to several thousands pounds

along with annual savings of several million

dollars in operating and acquisition costs.

Installation time is reduced, as the system only

requires the mechanical installation of the

actuator and the connection of two or more

electrical power cables. The requirement for

periodic maintenance is greatly reduced.

Additional building facilities are not required

for electromechanical actuators, whereas, a

hydraulic pump unit may require a separate

room with fluid containment provisions.

Engineers and actuation suppliers are looking

closer at the application of the roller screw as a

competing technology for the ball screw based

EMA. Traditionally it has been held that roller

screw is superior to ball screw technology.

Operational lifetime is often one of the biggest

attributes of the roller screw. Roller screws

have a higher dynamic load rating than ballscrews. Even though roller screws are reported

to be more costly than ball screws, the increased

load carrying capability of the roller screw may

make it an attractive technology for future EMA

applications.

When considering replacement of a hydraulic

actuator with an electric actuator, two basic

statements are appropriate: (1) a hydraulic

actuator will generally weigh less than an

electric actuator designed for the sameapplication, and (2) the hydraulic actuator will

usually be capable of delivering higher power

than required for its application.

The fact that a hydraulic actuator will usually be

capable of delivering higher power than

required for a particular application is a

consequence of hydraulic actuator area usually

being sized by the maximum hinge moment

yielding a corner horsepower (Hcp). The corner


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horsepower for a hydraulic actuator can be

described by the following relationship.

180121082.1

3 = RH

Hp m

Where: Hcp= Corner Horsepower

Hm = Hinge Moment (in-lb)R = Rate (deg/sec)1.82 x 10

-3= 1/550 ft-lb per sec

Maximum control surface deflection and hence

actuator stroke are based on surface

effectiveness characteristics at low speed, where

hinge moments are low. Maximum surface

rates are also set at some speed conditions.

Therefore, a situation where hydraulic actuator

area is set at one condition, stroke is set at

another and the rate at yet a third flight

condition. These results in a large-bore and

long-stoke actuator that must move at high rate.

This condition defines a unit with high horse

power capability that is never used [3].

The premise here is that an EMA with uniquecapabilities need not furnish anywhere near the

power of the hydraulic actuator it replaces. An

EMA only needs to have sufficient torque at one

point and adequate speed at another. Therefore,

horsepower capability is not a valid measure for

comparison.

3.0 Key Design Issues

Duty cycleis the most critical design parameter

for the electric actuator. It affects themechanical, thermal and electrical design of the

electric actuator. Mechanically, duty cycle will

predict the magnitude and frequency of loads.

This will impact the design to ensure that the

EMA meets the expected life requirements and

insure that the EMA has the ability to produce

the necessary force output. Duty cycle will

predict the rates and cycle data and also the

number of cycles per flight/maneuvers which

define the fatigue characteristics of the actuator.

The integration of EMA duty cycle with respect

to its expected life can be analyzed to determine

the total linear travel distance of the EMA. This

analysis is used to size the screw for its

expected L10 life and provides a better

understanding of the EMA fatiguecharacteristics [4]. The expected L10 life of a

roller screw or ball screw can easily be

computed and is expressed as the linear travel

distance that 90% of the screws are expected to

meet before experiencing metal fatigue.

6

3

10 10xF

CL

m

=

L10= life corresponding to a 10 percent probability

of screw failure

C = dynamic load rating of the roller screw and ball

screw nut assembly

This expression is then integrated with the cubic

mean load (Fm) with the EMA duty cycle and

the applied load (F1) and the distance (S1) the

screw travels with a particular load with respect

to time at load. The cubic load is computedusing the expression:

3

21

33

2

3

1

n

nm

SSS

FFFF

K

K

+

+=

Fm= mean load

F1 Fn= applied load

S1 Sn= distance of screw travel between loads

Thermally, duty cycle will help predict a

thermal load profile created by the actuator asdefined by the flight maneuvers/mission

profiles. Important factors like load duration,

load interval, rate of actuation, load, etc.

generate thermal energy that must be

considered. The components of the actuator

will need to be designed to handle these heat

loads. A thermal management system must be

sized for the actuator to mitigate the risk of

overheating the actuator. Duty cycle must be

analyzed from the standpoint of holding load

and the load with respect to expected life. Forholding loads are considered to determine how


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hard the EMA is working. This analysis is used

to size the motor torque capability and an

acceptable level of operating time so that the

thermal limits of the motor or actuatorcomponents are not exceeded which results in

tripping the thermal overload (at best) or burn

up the motor (at worst) [5]. The thermal duty

cycle is the ratio or percentage of actuator on

time to off time and is expressed as shown

below.

100x

TT

TD

offon

onc

+

=

Dc= duty cycle

Ton= time on

Toff= time off

Electrically, duty cycle serves as the input to an

actuator model to establish the power required

during flight, maneuvers, and etc. This help to

size the power management and generation

systems.

Jam-Resistant EMA technology needs to be

developed and demonstrated to ensure

widespread usage of electric actuation. If there

were no jamming concerns, the EMA can be the

simplest and most compact actuator. Jam-

resistant EMAs can be designed, but the system

may no longer be simple. Some jam-resistant

EMAs add other failure modes, which increase

the risk of failure. As the complexity goes up,

the inertia of the EMA goes up and the response

goes down. The higher the inertia, the higher

the starting torque required, and the more heat

generated. Generally, response goes down as

the actuators get larger. Boeing is working with

major suppliers and research institutions to

develop EMAs that are fault tolerant and jam

resistant [6].

Power Densities of current motor-drive

technologies need to exceed 1 kW/lb to meet the

need of future military, space and commercialplatforms. Studies have been conducted to

evaluate the electric motor power densities that

exceed 1kW/lb as well as a fault tolerant

architecture. A specific research focus has

been to evaluate the suitability of switchedreluctance (SR) motor technology for EMAs.

SR motor drives are considered to be robust and

fault tolerant. These are important attributes

given the hostile environment of an aircraft in

flight and the safety-critical nature of the

application [7].

Motor Selectionlike duty cycle can sometimes

be the most challenging aspect of designing and

sizing and electric actuation system. A priority

list must be made as to which properties of the

motor system will be optimized. These

properties may include motor efficiency, motor

torque, motor power, reliability, and of course,

cost. Generally, torque is the driving factor in

motor weight, size, and consequently, cost.

Therefore, knowing the torque requirements is

paramount.

The torque for an SRM machine can be

described by the following relationship:

( )dt

idLiTe

,

2

1 2 =

Where: Te= Air gap Torquei = current

( ),

,

dt

idL Variation of inductance

Social Change, though nota technology issue,

is one of the major hurdles for widespreadacceptance of EMA technology. Over the years,

there has been a reluctance to make a major leap

toward a widespread use of electric actuation,

especially EMAs. Actuator designers and air

framers have a well-documented database for

conventional hydraulic system technology. This

database is supported by well-developed

specifications, ground rules, and even rules of

thumb that have been developed over the past

50 years. A similar database does not exist for

electric actuators, especially EMAs. To


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achieve the desired weight benefits, hydraulic

system designers addressed the issue of reduced

weight by going to higher pressure hydraulic

systems and developing lighter weight plumbingand fitting systems.

4.0 Electrical Power and Distribution System

The electrical power and distribution system

distributes the electrical power from various

power sources to aircraft loads. The increased

electrical power required for the MEA may only

economically be achieved through utilization of

reasonable high voltages AC and DC

distribution systems. 270 Vdc appears to be the

current choice for current electrical actuation

systems. In some cases traditional 115 Vac and

28 Vdc has been used for compatibility reasons

with present aircraft electrical systems

infrastructure. In these cases power conversion

units (PCUs) can be installed to rectify the 3-

phase, 115 V ac aircraft supply into the 270 V

dc power required by the actuator [8].

High power, electric actuation systemsare being proposed on many new aircraft with

ratings up to 50 kW. The electrical loads

presented by these actuators are dynamic in

nature and appear highly nonlinear, drawing

non-sinusoidal current waveforms. This has a

significant effect on the quality of the power

system voltage as it introduces harmonic noise

thatis seen by other connected loads [8]. These

requirements are placing a demand for more

stringent power quality, and this means

increasing the number of solid-statecomponents. Many of these electric actuators

and power control units operate at high powers

during only certain portions (take-off, flight

maneuvers and landing) of the mission. The

temperatures associated with these systems must

be maintained at threshold values that will not

have a negative effect on the electronic and

electrical component reliability.

5.0 Integrated Thermal Management

Thermal Management is quickly becoming a

limiting design factor for electric actuationsystems for future military aircraft and

commercial airplanes. The design of the thermal

management system for the EMA is a function

of the heat rejection of the drive and the

operational duty cycle. Future cooling demands

will require an integrated thermal management

strategy at the aircraft, subsystem, and

component levels for MEA to become a reality.

A new approach of dealing with the individual

component heat loads at a local level was

perceived to best achieve the goals of the MEA

Initiative. There are a number of thermal

management approaches and technologies

including: thermal energy storage, advanced

liquid cooling provides greater integration

between power generation and thermal

management systems. Advanced materials such

as nano technology and high thermal

conductivity graphite foams are being

developed that offer potential near and far term

thermal management advantages. Oak RidgeNational Laboratory (ORNL) has developed

graphite foam (Figure 4) that has a specific

thermal conductivity six times more than copper

and five times more than aluminum [9].

Figure 4. Graphite Foam Heat Sink

These new technologies allow engineers to

evaluate concepts that will thermally integrate

and structurally embed the electric actuator and

its associated electronics into the vehicle

structure. This integration approach concept

allows the heat generated by the actuator to bemanaged by a passive system that could use


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existing airframe structure and other advanced

thermal management technology to transfer the

heat to a nearby heat sink. This passive cooling

system has the potential of offering the best nearterm solution in terms of reduced weight,

reduced maintenance and reduced operation

cost.

6.0 Integrated Health Management

7

Integrated Health Management for electric

actuation systems includes failure analysis of

the system and the components. It requires a

thorough understanding of the underlying root

causes of device failuresfailure physicsas a

basis for credible and reliable predictions on

remaining life of the device, component, or

equipment.

It has long been a sound engineering practice to

utilize S-N (stress-number of cycles) curve as a

mean to determine the expected life of a

component while operating at a level where the

stress, strain, or similar metric can be measuredand the cycles counted until the lower bound

was exceeded and the component would be

retired from further service. These components

were evaluated based on high-cycle fatigue

(HCF) >10,000 cycles and low cycle fatigue

(LCF)

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David Blanding

Electromechanical actuators are a cost-effective

alternative because there is one energy

conversion versus two in a hydraulic system.

Because of absence of hydraulic fluid,electromechanical actuators will be the

preferred solution for most MEA aerospace

platforms.

Advances in electric motor and power

electronics are making the technology both

affordable and cost effective as well as

providing a greater opportunity for expanded

subsystem integration.

Thermal management is a critical technology

need and more integrated solutions are required

to minimize the adding of additional weight to

cool the electric motor and power electronics.

Concepts for integrated thermal management

approaches are being evaluated within the

aerospace community.

Though not a technology hurdle, social change

is one of the major hurdles for widespread

acceptance of EMA technology.

8.0 Acknowledgement

The author wishes to acknowledge that the

content and many of the thoughts advanced

herein have evolved from papers, presentation

and discussions with many individuals. This

paper has been influenced by previous work

done by the author while employed at TheBoeing Company and by others in the aerospace

industry

9.0 References

1. Wallace J., Seattle Post-Intelligencer, May

20, 2004

2. Blanding D., Current Development in

Electric Actuation Technology, Boeing

BTECH 2004, June 2004

3. John B. Leonard, A System Look at

Actuation Concepts and Alternatives for

Primary Flight Control, SAE Paper

851753.

4. Hicks T., Handbook of Mechanical

Engineering Calculation, McGraw-Hill,

1998.

5. Blanding D., Grow J., Chen J., Evaluating

Electric Actuation Duty Cycle Thru the

Application of Confidence Interval, Eighth

Boeing Technical Excellence Conference,February 2004.

6. Blanding D., Tesar D., Krishnan R., Fault

Tolerant Electromechanical Actuation

Project Progress Report, AVS-1ME7K003,

Boeing Phantom Works, dated October 11,

2002.

7. Krishnan R., Blanding D., Bhano A., Ha K.,

and Morse J, High Reliability SRM Drive

System for Aerospace Applications, Center

for Rapid Transit Systems, Virginia Tech.

8. Karimi K., Electric Power System Design,Power Quality, Boeing MEA Seminar

Series, December 2005.

9. Klett J., High Thermal Conductivity

Mesophase Pitched-Derived Carbon

Foams, Oak Ridge National Laboratory,

Oak Ridge Tennessee.

10.Annis, C., Random Fatigue Limit on a P/C,

Staticistical Engineering, January 2006


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